Combustor for a gas turbine

ABSTRACT

A combustor for a gas turbine having: a pre-combustion chamber, a swirler, a pilot burner upstream the pre-combustion chamber which has a pilot burner surface separating the pilot burner from the pre-combustion chamber. The pilot burner further includes at least a pilot fuel injector, wherein the combustor includes a lip extending from the pilot burner surface in the pre-combustion chamber, the lip including an internal surface oriented towards the pilot fuel injector, the internal surface being inclined of an inclination angle of between 0 degrees and 90 degrees with respect to the center axis of the pre-combustion chamber, and the lip has at least a feed passage for connecting the internal surface with the flow of oxidant gas coming from the swirler.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is the US National Stage of International ApplicationNo. PCT/EP2016/080488 filed Dec. 9, 2016, and claims the benefitthereof. The International Application claims the benefit of EuropeanApplication No. EP15202500 filed Dec. 23, 2015. All of the applicationsare incorporated by reference herein in their entirety.

FIELD OF INVENTION

The present invention relates to a combustor for a gas turbine.

ART BACKGROUND

In such a technical field, a combustor generally comprises a maincombustion chamber and a pre-combustion chamber, upstream of the maincombustion chamber. The pre-combustion chamber comprises a swirlersection having a swirler through which a main fuel stream is provided.In the swirler the main fuel is mixed to a non-combustible gas flowcomprising an oxidant, for example air. The main fuel stream and thenon-combustible gas flow are injected via the swirler into thepre-combustion chamber of the combustor in a generally tangentialdirection with respect to the centre axis of the combustor.

A pilot fuel is further injected in the pre-combustion chamber forcontrolling the combustor flame in which the main fuel in burned. Thepilot fuel is typically injected by a pilot burner, generally accordinga direction parallel to the centre axis of the combustor.

The pilot fuel is injected from the pilot burner into the pre-combustionchamber through a plurality of pilot fuel injectors arranged on thepilot burner surface, i.e. the surface separating the pilot burner fromthe pre-combustion chamber. The main fuel and the pilot fuel is agaseous fuel. Liquid fuel injection may also be provided in similarpositions on the swirler and on the pilot burner.

The combustion of the pilot fuel is achieved through an oxidant, forexample air, first being mixed together with the fuel in the pilotburner.

In known solution, the injected pilot fuel generates a diffusion flameinside the pre-combustion chamber, close to pilot burner surface. Thishas the main drawback of increasing the local temperature at the pilotburner surface, with the consequence of reducing the life cycle of thepilot burner.

It is therefore desirable to provide a new design of the combustor abovedescribed, in particular at the interface between the pilot burner andthe pre-combustion chamber, for limiting temperatures at the pilotburner surface, at the same time without compromising the overallefficiency of the combustor. Inside the combustor, avoiding areas withhigh temperature has also the positive effect in reducing overallnitrogen oxides (NOx) emissions.

SUMMARY OF THE INVENTION

It may be an objective of the present invention to provide a combustorsolving the above described inconveniences experimented in knowncombustors.

It may be a further objective of the present invention to provide acombustor with a proper fuel distribution in the mixture of the gasinside the pre-combustion chamber, in order to avoid areas withnon-desirable high temperature.

It may be another objective of the present invention to provide acombustion chamber with an improved life-cycle of components subject tohigh temperature, in particular the pilot burner.

This object is solved by a combustor for a gas turbine according to theindependent claim. The dependent claims describe advantageousdevelopments and modifications of the invention.

According to an aspect of the present invention, a combustor for a gasturbine is presented. The combustor comprises: a pre-combustion chamber,a swirler which is connected to the pre-combustion chamber for providingpre-combustion chamber with a flow of oxidant gas. The swirler isarranged around the pre-combustion chamber in a circumferentialdirection with respect to an axis of the pre-combustion chamber, a pilotburner upstream the pre-combustion chamber which comprises a pilotburner surface separating the pilot burner from the pre-combustionchamber. The pilot burner further comprises at least a pilot fuelinjector which is arranged to the pilot burner surface for injectingpilot fuel into the pre-combustion chamber.

The combustor includes a lip extending from the pilot burner surface inthe pre-combustion chamber, the lip including an internal surfaceoriented towards the pilot fuel injector for intercepting at least partof the pilot fuel from the pilot fuel injector, the internal surfacebeing inclined of an inclination angle with respect to the centre axisof the pre-combustion chamber, the inclination angle being comprisedbetween 0 degrees and 90 degrees. The lip comprises one or more feedpassage for connecting the internal surface with the flow of oxidant gascoming from the swirler.

The combustor may be an annular-type or a can-type combustor. Thecombustion chamber may have a cylindrical or oval shape. The combustionchamber may comprise a main combustion chamber and a pre-combustionchamber with a swirler section. The centre axis of the pre-combustionchamber may be a symmetry line of the pre-combustion chamber. At theswirler section, the swirler is mounted to the pre-combustion chamberand surrounds the pre-combustion chamber centre axis.

Advantageously, the inclined orientation of the lip guides the flow awayfrom the pilot burner surface and towards a main combustion zone of thepre-combustion chamber. The injection of oxidant gas through the feedpassages enhance mixing of the oxidant gas with the pilot fuel from thepilot fuel injector. As a result, temperature at the pilot burnersurface is reduced, up to more acceptable values, which make life of thepilot burner longer.

According to possible embodiments, an inclination angle is comprisedbetween 30° and 60° has proved to be particularly advantageous.

According to possible embodiments of the present invention, the lipfurther comprises an external surface oriented towards the swirler forintercepting at least part of the flow of oxidant gas coming from theswirler, the feed passages being provided between the internal surfaceand the external surface. The feed passages may be provided inplurality, regularly distributed around the centre axis.

The external surface may be provided with a plurality of turbulators forinducing turbulence in at least part of the flow of oxidant gas comingfrom the swirler. The turbulators may comprise a plurality ofprotrusions extending orthogonally from the external surface and/or aplurality of channels having a depth extending from the external surfacetowards the internal surface. The protrusions may comprise a circularrim concentric with the centre axis of the pre-combustion chamber.

Advantageously, the turbulators enhance the turbulence from the swirleroxidant gas to mix with the pilot fuel emerging from the lip towards theinside of the pre-combustion chamber. This produces premixed pilot forlowering temperatures and hence NOx emissions.

According to possible embodiments of the present invention, the internalsurface and the external surface have a common trailing edge, at the endof the lip, where both the pilot fuel and the oxidant gas separate fromthe lip. The trailing edge may have a circular profile around the centreaxis of the pre-combustion chamber. The trailing edge may have a wavedprofile.

Advantageously, the above described designs of the end portion of thelip improve turbulence and flow aerodynamics. Pressure loss may be alsoreduced.

According to other embodiments of the present invention, the internalsurface has an aerofoil shape. Advantageously, this improves turbulenceand flow aerodynamics and reduces pressure loss.

According to further embodiments of the present invention, the lip isprovided as an edge of a shroud of the pilot burner extending inside thepre-combustion chamber. Advantageously, this allows to manufacture apilot burner directly including a lip optimised for the presentinvention.

BRIEF DESCRIPTION OF THE DRAWINGS

The aspects defined above and further aspects of the present inventionare apparent from the examples of embodiment to be described hereinafterand are explained with reference to the examples of embodiment. Theinvention will be described in more detail hereinafter with reference toexamples of embodiment but to which the invention is not limited.

FIG. 1 shows a longitudinal sectional view of a gas turbine engineincluding a combustor according to the present invention,

FIG. 2 shows a partial and schematic longitudinal section of a combustorfor a gas turbine according to an exemplary embodiment of the presentinvention, showing a pilot burner, a pre-combustion chamber and aswirler section;

FIG. 3 shows a sectional view of a swirler according to exemplaryembodiments of the present invention, according to the section lineIII-III of FIG. 2;

FIG. 4 shows a magnified view of the detail IV of FIG. 2;

FIG. 5 shows an assonometric partial view of the combustor for a gasturbine, according to an exemplary embodiment of the present invention,partially showing a pilot burner;

FIG. 6 shows a partial sectional view of the combustor of FIG. 5;

FIG. 7 shows a partial sectional view, corresponding to the sectionalview of FIG. 6, of another embodiment of a combustor for a gas turbineaccording to the present invention;

FIGS. 8 to 10 show three assonometric partial views, corresponding tothe assonometric view of FIG. 5, of other respective embodiments of acombustor for a gas turbine according to the present invention.

DETAILED DESCRIPTION

The illustrations in the drawings are schematic. It is noted that indifferent figures, similar or identical elements are provided with thesame reference signs.

FIG. 1 shows an example of a gas turbine engine 10 in a sectional view.The gas turbine engine 10 comprises, in flow series, an inlet 12, acompressor section 14, a burner section 16 and a turbine section 18which are generally arranged in flow series and generally about and inthe direction of a longitudinal or rotational axis 20. The gas turbineengine 10 further comprises a shaft 22 which is rotatable about therotational axis 20 and which extends longitudinally through the gasturbine engine 10. The shaft 22 drivingly connects the turbine section18 to the compressor section 14.

In operation of the gas turbine engine 10, air 24, which is taken inthrough the air inlet 12 is compressed by the compressor section 14 anddelivered to the combustion section or burner section 16.

The burner section 16 comprises a burner plenum 26, one or morecombustion chambers 28, each having a respective upstream pre-combustionchamber 101. The burner section 16 further comprises at least one pilotburner 30 and a swirler section 31 fixed to each pre-combustion chamber101. The pre-combustion chambers 101, the combustion chambers 28, thepilot burners 30 and the swirler section 31 are located inside theburner plenum 26. The compressed air passing through the compressor 14enters a diffuser 32 and is discharged from the diffuser 32 into theburner plenum 26 from where a portion of the air enters the pilot burner30 and is mixed with a gaseous or liquid pilot fuel. The air/fuelmixture is then burned and the combustion gas 34 or working gas from thecombustion is channelled through the combustion chamber 28 to theturbine section 18 via a transition duct 17.

A main flow of air/fuel mixture is further inserted in thepre-combustion chamber 101 through the swirler section 31, as betterdetailed in a following section of the present text. The main fuel burnswhen mixing with the hot gasses in the pre-combustion chamber 101 and inthe main combustor chamber 28.

This exemplary gas turbine engine 10 has a cannular combustor sectionarrangement, which is constituted by an annular array of combustor cans19 each having a pilot burner 30 and a combustion chamber 28, thetransition duct 17 having a generally circular inlet that interfaceswith the combustor chamber 28 and an outlet in the form of an annularsegment. An annular array of transition duct outlets form an annulus forchannelling the combustion gases to the turbine 18.

The turbine section 18 comprises a number of blade carrying discs 36attached to the shaft 22. In the present example, two discs 36 eachcarry an annular array of turbine blades 38. However, the number ofblade carrying discs could be different, i.e. only one disc or more thantwo discs. In addition, guiding vanes 40, which are fixed to a stator 42of the gas turbine engine 10, are disposed between the stages of annulararrays of turbine blades 38. Between the exit of the combustion chamber28 and the leading turbine blades 38 inlet guiding vanes 44 are providedand turn the flow of working gas onto the turbine blades 38.

The combustion gas from the combustion chamber 28 enters the turbinesection 18 and drives the turbine blades 38 which in turn rotate theshaft 22. The guiding vanes 40, 44 serve to optimise the angle of thecombustion or working gas on the turbine blades 38.

The turbine section 18 drives the compressor section 14. The compressorsection 14 comprises an axial series of vane stages 46 and rotor bladestages 48. The rotor blade stages 48 comprise a rotor disc supporting anannular array of blades. The compressor section 14 also comprises acasing 50 that surrounds the rotor stages and supports the vane stages48. The guide vane stages include an annular array of radially extendingvanes that are mounted to the casing 50. The vanes are provided topresent gas flow at an optimal angle for the blades at a given engineoperational point. Some of the guide vane stages have variable vanes,where the angle of the vanes, about their own longitudinal axis, can beadjusted for angle according to air flow characteristics that can occurat different engine operations conditions.

The casing 50 defines a radially outer surface 52 of the passage 56 ofthe compressor 14. A radially inner surface 54 of the passage 56 is atleast partly defined by a rotor drum 53 of the rotor which is partlydefined by the annular array of blades 48.

The present invention is described with reference to the above exemplaryturbine engine having a single shaft or spool connecting a single,multi-stage compressor and a single, one or more stage turbine. However,it should be appreciated that the present invention is equallyapplicable to two or three shaft engines and which can be used forindustrial, aero or marine applications.

The terms upstream and downstream refer to the flow direction of theairflow and/or working gas flow through the engine unless otherwisestated. When not differently specified, the terms axial, radial andcircumferential are made with reference to an axis 35 of the combustor.

FIG. 2 shows a combustor 100 for a gas turbine. The combustor 100 has acentre axis 35 and comprises:—an upstream portion with a pre-combustionchamber 101 and a swirler 103, and—a downstream portion with acombustion chamber 28.

The pre-combustion chamber 101, the swirler 103 and the combustionchamber 28 are all axially symmetric around the centre axis 35. Withrespect to the centre axis 35, the pre-combustion chamber 101 has asmaller diameter than the combustion chamber 28. The pre-combustionchamber 101 and the combustion chamber 28 are adjacent to one anotheralong the centre axis 35 and in fluid communication with one another.Downstream of the pre-combustion chamber 101 the combustion chamber 28extends up to the transition duct 17. The combustion chamber 28 isconventional and therefore not described in further detail.

The swirler 103 is mounted on a peripheral wall 115 of thepre-combustion chamber 101, in such a way that the swirler 103 surroundsthe pre-combustion chamber 101 in a circumferential direction withrespect to the centre axis 35. The swirler 103 comprises a bottomsurface 104 which is orthogonal to the centre axis 35 and which forms apart of a slot 201 (see FIG. 3) through which, typically, anoxidant/fuel mixture flow F is injectable into the pre-combustionchamber 101.

The swirler 103 further comprises a cylindrical peripheral surface 119having axis coincident with the combustor centre axis 35,

With reference to FIG. 3, the swirler 103 comprises a plurality of slots201 (twelve slots in the embodiment of FIG. 3). Each slot 201 is formedby circumferentially spaced apart vanes 203 and the bottom surface 104.Oxidant/fuel mixture which flows through the slots 201 is directedapproximately tangentially with respect to the centre axis 35. Thisorientation of the slots 201 induces a swirl movement, i.e. a movementaccording to a tangentially orientated direction around the centre axis35, of the gasses inside the pre-combustion chamber 101.

Each slot 201 comprises a base fuel injector 107 which is arranged tothe bottom surface 104 such that an air/fuel mixture is injectable intothe slot 201 according to a main fuel injection direction which isorthogonal or inclined with respect to the bottom surface 104.

Additionally, further side fuel injectors 202 may be provided for someof the slots 201 or for all of the slots 201 on the cylindricalperipheral surface 119 of the swirler 103.

In the embodiment of the attached figures two side fuel injectors 202are provided for each of the slots 201.

The side fuel injectors 202 inject further fuel. The further fuel may bemixed inside the slots 201 with the fuel which is injected by the basefuel injector 107 and with the oxidant.

Side fuel injectors 202 are in the form of holes, injecting furthergaseous fuel.

According to other embodiments of the present invention, atomizers ornozzles for liquid fuel injection are provided in the same slots 201,close to the trailing edges of the swirler vanes 203.

Upstream to the swirler 103 and to the pre-combustion chamber 101, thecombustor 100 further comprises a pilot burner 110, which comprises aburner face 111. In particular, the burner face 111 is aligned orsubstantially parallel to the bottom surface 104. The pilot burner 110further comprises a cylindrical shroud 170, extending around the centreaxis 35, for peripherally delimiting the pilot burner 110.

The pilot burner 110 comprises a plurality of pilot fuel injectors 112which are arranged to the burner face 111 for injecting pilot fuel intothe pre-combustion chamber 101. In the embodiments of the attachedfigures, twelve side pilot fuel injectors 112 regularly distributed 30degrees apart circumferentially around the centre axis 35 are provided.

The pilot fuel injectors 112 are oriented substantially parallel to thecentre axis 35.

The pilot fuel forms a separation layer and a flame front 105. Thecirculation induced by the radial swirler 103 forms a central circularzone around the centre axis 35, inside of which the pilot fuel (i.e. theoxidant/fuel mixture) is burned. This central zone is called thereaction zone RZ. Around the central reaction zone RZ, the oxidant/fuelmixture is injected by the swirler 103.

With reference to FIGS. 4 to 10, the combustor 100 further includes alip 150 extending from the pilot burner surface 111 in thepre-combustion chamber 101. In a circumferential direction, the lip 150further extends around the centre axis 35.

The lip 150 extends from a portion of the pilot burner surface 111 whosedistance from the centre axis 35 of the pre-combustion chamber 101 isgreater than the distance between the pilot fuel injectors 112 and thecentre axis 35.

With respect to the more internal portion of the pre-combustion chamber101, identified as the portion around the centre axis 35, the lip 150includes an internal surface 151 and an external surface 152.

The internal surface 151 is inclined towards the centre axis 35 andoriented towards the pilot fuel injectors 112 for intercepting at leastpart of the pilot fuel from the pilot fuel injectors 112. With respectto the centre axis 35, the internal surface 151 is inclined of aninclination angle α comprised between 0 degrees and 90 degrees. Moreparticularly, in the embodiments of FIGS. 4 to 10, the inclination angleα is comprised between 30 degrees and 60 degrees. The external surface152 is oriented towards the swirler 103 for intercepting at least partof the flow F coming from the swirler 103.

The lip 150 is integral with the pilot burner 110, being provided as anedge of the shroud 170, extending inside the pre-combustion 101.

According to other embodiments of the present invention (not shown) thelip 150 is provided on pilot burner surface 111 or on the swirler 103.

The lip 150 further comprises a plurality of feed passages 155 providedbetween the internal surface 151 and the external surface 152, forconnecting the internal surface 151 with the flow F coming from theswirler 103. The feed passages 155 are regularly distributed around thecentre axis 35.

With specific reference to FIGS. 8 and 10, the external surface 152comprises a plurality of turbulators 160, 161, 162 for inducingturbulence in the flow F coming from the swirler 103.

In the embodiment of FIG. 8, the turbulators comprise a plurality ofprotrusions 160, 162 extending orthogonally from the external surface152. Some of the protrusions 160, 162 are constituted by a plurality offirst protrusions 160, placed around the centre axis 35, at a samedistance from the centre axis 35. The first protrusions 160 haverespective bases on the external surface 152, the bases having, forexample, circular or rectangular shape. The protrusions 160 areregularly distributed around the centre axis 35, at a fixed angulardistance. A further protrusion 162 is provided as a circular rim 162,concentric with the centre axis 35 of the pre-combustion chamber 101.With respect to the flow F coming from the swirler 103, the circular rim162 is provided on the external surface 152, downstream of the firstprotrusions 160. According to other possible embodiments (not shown),the circular rim 162 is provided on the external surface 152, upstreamof the first protrusions 160.

In the embodiment of FIG. 10, the turbulators comprise a plurality ofchannels 161 regularly distributed around the central axis Y. Eachchannel 161 extends from the external surface 152 up the internalsurface 151, in such a way that the channels 161 divide the lip 150 intoa plurality of segments 158, each segment 158 being comprised betweentwo consecutive channels 161. According to other possible embodiments(not shown), the channels 161 are not completely extended from theexternal surface 152 up the internal surface 151, but are provided onthe external surface 152 along a direction inclined of an inclinationangle α with respect to the centre axis 35 and with a depth extendingfrom the external surface 152 towards the internal surface 151.

In other embodiments of the present invention (not shown) othercombination of the turbulators 160, 161, 162 described above may bepossible. In particular, any other array of the turbulators 160, 161,162 may be arranged, each array being characterized by the type(s),number and distribution of the turbulators 160, 161, 162.

The internal surface 151 and the external surface 152 have a commontrailing edge 156, at the end of the lip 150, where both the pilot fueland the flow F separate from the lip 150.

With specific reference to FIGS. 5, 6, and 8, the trailing edge 156 hasa sharp circular profile around the centre axis 35. With specificreference to FIG. 9, the trailing edge 156 has a rounded profile in asection view (equivalent, for example to the view of FIG. 6) and a wavedprofile in a circumferential view, around the centre axis 35.

With reference to FIGS. 7 and 10, the trailing edge 156 is clipped, i.e.the lip 150 has, in a sectional plane including the centre axis 35, atrapezoidal shape, including a face 159, connecting the external surface152 and the internal surface 151, at the end of the lip 150.

With further reference to FIG. 6, the lip 150 may be provided, inembodiments of the present invention, with an internal surface 151 bhaving an aerofoil shape (dashed line of FIG. 6).

It should be noted that the term “comprising” does not exclude otherelements or steps and “a” or “an” does not exclude a plurality. Alsoelements described in association with different embodiments may becombined. It should also be noted that reference signs in the claimsshould not be construed as limiting the scope of the claims.

1. A combustor for a gas turbine, comprising: a pre-combustion chamber,a swirler which is connected to the pre-combustion chamber providingpre-combustion chamber with a flow of fuel and oxidant gas, the swirlerbeing arranged around the pre-combustion chamber in a circumferentialdirection with respect to a centre axis of the pre-combustion chamber, apilot burner upstream the pre-combustion chamber which comprises a pilotburner surface separating the pilot burner from the pre-combustionchamber, the pilot burner further comprising at least a pilot fuelinjector which is arranged to the pilot burner surface for injectingpilot fuel into the pre-combustion chamber, wherein the combustorincludes a lip extending from the pilot burner surface in thepre-combustion chamber, the lip including an internal surface orientedtowards the pilot fuel injector for intercepting at least part of thepilot fuel from the pilot fuel injector, the internal surface beinginclined of an inclination angle (α) with respect to the centre axis ofthe pre-combustion chamber, the inclination angle (α) being comprisedbetween 0 degrees and 90 degrees, and wherein the lip comprises at leasta feed passage connecting the internal surface with the flow of oxidantgas coming from the swirler.
 2. The combustor according to claim 1,wherein the lip further comprises an external surface oriented towardsthe swirler intercepting at least part of the flow coming from theswirler, the feed passage being provided between the internal surfaceand the external surface.
 3. The combustor according to claim 2, whereinthe external surface comprises a plurality of turbulators inducingturbulence in at least part of the flow of oxidant gas coming from theswirler.
 4. The combustor according to claim 3, wherein the turbulatorscomprise a plurality of protrusions extending orthogonally from theexternal surface.
 5. The combustor according to claim 3, wherein theprotrusions comprise a circular rim, the circular rim being concentricwith the centre axis of the pre-combustion chamber.
 6. The combustoraccording to claim 3, wherein the turbulators comprise a plurality ofchannels having a depth extending from the external surface towards theinternal surface.
 7. The combustor according to claim 2, whereininternal surface and the external surface have a common trailing edge.8. The combustor according to claim 7, wherein the trailing edge has acircular profile around the centre axis of the pre-combustion chamber.9. The combustor according to claim 7, wherein the trailing edge has awaved profile.
 10. The combustor according to claim 1, wherein theinternal surface has an aerofoil shape.
 11. The combustor according toclaim 1, wherein the value of the inclination angle (α) is comprisedbetween 30° and 60°.
 12. The combustor according to claim 1, wherein thelip extends from a portion of the pilot burner surface whose distancefrom the centre axis of the pre-combustion chamber is greater than thedistance between the pilot fuel injector and the centre axis.
 13. Thecombustor according to claim 1, wherein the lip is provided as an edgeof a shroud of the pilot burner extending inside the pre-combustionchamber.
 14. The combustor according to claim 1, wherein the lipcomprises a plurality of feed passages connecting the internal surfacewith the flow of oxidant gas coming from the swirler, the feed passagesbeing regularly distributed around the centre axis.